Rotative winged aircraft



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Jan. l, 1952l H.s. CAMPBELL ROTATIVE WINGED AIRCRAFT 5 Sheets-Sheet l Filed July 17, 1945 ATTOR EY$ Janl, 1952 H.-s. CAMPBELL 2,580,514

ROTATI-VE wINGED AIRCRAFT Filed July 17, 1945 5 SheetS-Sheet 2 IQNENTOR I @www TTORNEYJ Jan. 1, 1952 H. s. CAMPBELL ROTATIVE WINGED AIRCRAFT 5 Sheets-Sheet 3 Filed July 17, 1945 vw @E NN m 5 Sheets-Sheet 4 H. S. CAMPBELL ROTATIVE WINGED AIRCRAFT Jan. 1, 1952 Filed July 17, `1945 Jan. l, 1952 H. s.- CAMPBELL ROTATIVE WINGED AIRCRAFT,

Filed July 17. 1945 5 Sheets-Sheet 5 NEA/v F170/ CYCLC FITC/l Q:WENTOR BY A M ATTORNEYJ Patented Jan. 1, 19.52

ROTATIV WINGED AIRCRAFT Harris S. Campbell, Bryn Athyn, Pa., asllgnor to Autogiro Company of America, Philadelphia, Pa., a corporation of Delaware Application July 17, 1945, Serial No. 605.577

Claims.

r'his invention relates to rotative winged air-` craft and particularly to improvements in the hub or head structure of an aircraft sustaining rotor. Various features of the invention are usefui in rotative winged aircraft of a variety of types having one or more sustaining rotors, either nors mally power driven or normally autorotative. As

' ture incorporating an improved form of blade mounting including a floating hub member, and further incorporating blade pitch control means arranged to provide automatic stability with respect to floating movement of the'hub, so that in a given condition of operation, if a wind g-ust or other external force tends'to disturb the plane of rotation of the floating hub, the pitch angle of the rotor blades is v ried in a' manner to restore the hub to its orig nal plane of rotation.

Another object of the invention is the provision of an improved pitch control system providing for cyclical pitch variation as well as for alteration of the mean or average pitch of the rotor, various individual features and advantages of which will be pointed out more fully hereinafter.

In accordance with another aspect of the invention, provisionis made for restraining or locking out floating movements of the floating hub member when the aircraft is on the ground. In accomplishing this the invention provides a. releasable floating hub lock which is controlled in common with the mean pitch of the rotor in such a way that when the mean pitch is reduced to a non-lifting or to a substantially zero value, theV lock is automatically applied, the lock being automatically released upon actuation of the pitch commi to raise the mean blade citen-for instance when effecting vertical take-oil'.

Still another object of the invention is the arrangement of individual flapping and drag pivots for each blade in such manner that the drag pivot for each blade is operatively interposed be- (Cl. 17o-160.26)

ranged to incorporate limiting stops for both apping and lag-lead blade movements.. In association with the configuration of blade mounting pivots and parts, as Just mentioned, the invention further contemplates employment of `blade movement control devices, such as dampers for controlling movement of the blades about their drag pivots. By operatively interposing the drag pivots between the flapping pivots and the hub member, the' blade dampers or the like may readily be arranged to react directlybetween the s-everal blades of the rotor without the complication required to accommodate individual flapping movements of the blades.

In accordance with still another aspect of the invention, the rotor hub, power transmission including reduction gearing, clutches and power take-off shaft for driving the anti-torque airscrew constitute an assemly arranged .for unitary mounting in the body of the aircraft; the engine, together with a'cooling fan therefor, constituting another assembly independently mounted in the body of the aircraft, the two assemblies being adapted for ready coupling to each other.

How the foregoing together with other objects and advantages are attained will appear more fully from the following description referring to the accompanying drawings in' which- Figure 1 is an outline side view of a single rotor helicopter having a torque correcting iairscrew at the tail, the aircraft being constructed in accordance with the various features of the invention;

Figure 2 is a plan outline View of the aircraft shown in Figure 1; l

Figure 3 is a vertical sectional view. to an enlarged scale, of the rotor headand certain associated parts, this view being taken as indicated by the section line 3--3 on Figure 4;

Figure 4 is a plan view, partly in elevation and partly in horizontal section of the rotor head shown in Figure 3;

Figure 5 is a. horizontal sectional view taken as indicated by the section line 5-5 on Figure 3;

Figure 6 is a horizontal sectional view taken as indicated by section line 6-6 on Figure 3;

Figure 'I is a view showing portions of the floating hub mechanism, the view being taken substantially as indicated by line 1--1 on Figure 3: and

Figure 8 is an outline view of the rotor hub and power transmission assembly, with a somewhat diagrammatic illustration of the controls associated therewith.

In considering the drawings, attention is ilrst directed to Figures Land 2 in which the inven? tion is shown as applied to a helicopter type aircraft having a single generally centralized sustaining rotor incorporating three sustaining rotor blades Il and also having a torque counteracting tail rotor incorporating a pair o! blades il. 'I'he rotor and airscrew are -both adapted to be driven from the engine I2 which is mounted in the body I8 on a vertical axis, the engine being provided with a cooling fan -l4 mounted on the engine shaft i8 and adapted to circulate air downwardly over the engine, an air intake opening being provided at I8, just below the rotor head, and air outlets at I1, in the bottom of the fuselage. Certain features of the arrangement Just described are discosed and claimed in my copending application, Serial No. 702,523, filed October 10, 1946.

Thevoccupants compartment is arranged forward of the engine compartment and, in the particular vaircraft shown. is provided with pilots seats |8-I8, and an occupants seat I 8.

FromFigureslandBitwillbeseenthatthe rotor hub and power transmission assembly is provided with a downwardly extending shaft 28 adapted to be coupled with the engine shaft i8, through a short shaft 2l and a pair of universal Joints 22. Shaft 26 serves to deliver power from the engine into the casing 23 which incorporates a manually disconnectlble clutch and also an overrunning clutch providing for free rotation of the rotor and of the anti-torque airscrew in the event of engine failure. A second transmission casing appears at 24 which incorporates reduction gearing for the rotor drive. as is described more in detail herebelow with reference to Figures 3 and 6. The two casings 2l and 24 are interconnected by a casing sleeve 28 from which the power take-oil! shaft 28 extends, the shaft 28 being adapted to be coupled through shafting 21 with the anti-torque airscrew to drive the same.

With the manualI clutch and the overrunning clutch disposed in the power transmission line ahead of the power ltaire-oi! to the anti-torque rotor, it will be understood that the anti-torque rotor will always rotate with the sustaining rotor. It is contemplated that control of the aircraft in yaw be secured by variation of the mean pitch of the anti-torque air-screw, this control being effective either in power driven operation of the sustaining rotor or in free or autorotative operation thereof. The details of such control for the anti-torque airscrew need not be considered herein as they form no part of the present invention per se.

According to the invention the several transmission elements, together with the rotor head, comprise an assembly which is unitarily mounted in the body of the aircraft as by means of a centrally apertured conical support' 28 surrounding the transmission assembly, andlto which the upper casing part 24 is bolted. Removal of the fastening bolts and disconnection of ,shafting 2| and 21 permits lifting the entire transmission assembly (including the rotor hub) upwardly from the body, the diameter of the air intake 4 extended upwardly through the casing sleeve 28 into the gear casing 24. Shaft 3i is connected with the rotor drive spindle 82 by meansvr of epicyclic gearing. including a central pinion 8l, and a plurality of planet gears 34 mounted by means of a cage 85, the cage being splined to the rotor drive spindle 32 .and the planet gears 84 meshing with the internal ring gear 86 which, in eifect. forms a part of the gear casing 24. In this way a substantial reduction is effected in the drive to the rotor at a point beyond the power take-off to the anti-torque airscrew. The rotor drive spindle 82 which serves also to support the rotor is mounted by means of bearings l'l--i'l within a non-rotative supporting sleeve 88 which is rigidly associated with the gear casing 24. The thrust of lift or sustention of the rotor is therefore transmitted from the spindle 82 through the bearings 81 to the sleeve 28 and thence through gear casing 24 to the support 28 and the structural elements of the fuselage (see Figure 1).

At its upper end the spindle 32 is forked as at 32a-82a, the forks being apertured to receive a pivot structure composed of pivot parts 38-38 surrounding the ends obthe retainer pin 88a, the parts "-88 being supported in the'universal block 4l. As best seen in Figure 4, a pivot 4I, through which the retainer pin 88a passes,`

isalso mounted in the universal block 40 and extends into apertures formed in ears 42-42 which 'dependfrom the top closure member 48 of the floating hub member 44. `'.'l'lfie universal Joint (32a, 38, 4'0, 4| and 42) provides for free tilting movement of the floating hub member to various different positions, so that the hub member is therefore free to rotate' in various different planes.

In considering the attachment of the blades of the rotorfto the floating hub member, it is first noted that the rotor may incorporate a total number of blades different from that illustrated. the particular embodiment of the drawings incorporating three blades. From the standpoint of smoothness of operation of the rotor. particularly when incorporating a freely floating hub,

three blades are preferred, but it will be understood that variousvfeatures of the invention. including for example, vfeatures of the individual blade mounting pivots, floating hub locking mechanism, pitch control, etc. are also applicable to rotors incorporating a different number of blades. for example two blades or four blades. Moreover, various features of the blade mounti118, pitch control and of other structures are applicable to rotors having fixed spindle hubs, as weil as to rotors having floating hubs;

In the illustrated embodiment, with three blades, the floating hub member 44' is provided with three projecting blade mounting stubs 48 equally spaced angularly about the hub. Each blade is connected with one of the stubs 45 by means of individual pivots which are now described with reference to the blade at the rightopening i8 and the diameter of the central aperture in the support 28 being large enough to pass the upper and lower transmission casings 23 and 24.

Apertuies 28 in the conical wall of the support 28 serve to permit free circulation of air downwardly from the intake I6 to the engine i2.

As best seen in Figures 3 and 6, the power drive to the rotor includes a shaft 3| which is hand side lof Figures 3 and 4. The root end 48 of the blade is enlarged inwardly to form the external part 41 of a pitch mounting. This part is mounted by bearings 48 and 49 on a spindle 50, f

, s y whichis iournalled bybearings N-M in aver-l 'tured ears lI-ll formed at the outer end oi' the blade mounting stub 4l. The intersecting napping and drag pivots provide freedom for swinging movement of the blade in the napping sense (as indicated by the lines f-f in Figure 3) and also freedom for lag-lead movement (as indicated by lines l-l in Figure 4). Limiting stops i'or these blade movements are associated with the intersecting pivots and include a stop member Il projecting into the interior `iii the spindle Il and serving to denne the range ot ilapping movement f-f. stop l1 (see particularly Figure 4) projects radially inwardly toward the hub within the hollow stub 4l and serves to deilne the limit of lag-lead motion in the lagging sense. Stop 5l limits swinging of the blade `on the drag pivot in the leadingisense, they stops Il and Il.

being arranged to define the fange of lag-lead movement above referred to.

It will be understood that the blade swinging movements described just above take place with reference to tlne'noating'` hub member and thus that the napping and lag-lead movements are movements in addition to the I reedom o! motion provided by virtue or the universal or tilting mounting of the iloating hub member. In Figure 3 a typical range of tilting movement of the hub is indicated by the lines t-t.

As seen in Figures 8 ,and 4, a double-ended lever Bl-Bl is mounted on and splined to each of the drag pivot pins B3. Adjacent arms I! ot adjacent blades are linked together by means of a blade movement control device advantageously in the form of a pair of relatively telescoping elements Bil-II. Resistance to relative movement of the elements 6,0 and Bl may be provided either hydraulically or` by friction, las desired. The specific structure of the blade damper itsel! Ais not a part of the present invention per se although it is contemplated that the type of resistive force employed be non-rebounding.

In considering the blade damper arrangement it should be noted that the direct interconnection v of the actuating arms 59 around the hub pro-f vides for damping lag-lead movements ot the blades with relation to each other, without, howeve'r, appreciably resisting conjoint lag or lead movement of all blades in the same sense to the same degree. Moreover, since the drag pivots for the blades are operatively interposed between the napping pivots and the hub, and since the dampers operate through levers I! connected with the drag pivots, simple pivot connections are all that is needed between the damper-s and the levers M. Certain features oi' the blade movement control arrangement just described are disclosed and claimed in my copending application. Serial No. 702.524ii1ed October 10, i946.

Y For purposes of' control of the rotor blade pitch each blade is provided with a control arm $2 which is secured to the external part 41 or the pitch mounting and projects inwardly and forwardly therefrom, with respect to the direction o1' rotation of the rotor (see arrow 1- in Figure isaseociated. onesuehappearinginlligure 3. and the two being aligned 'along'an axis transverse to the axis ot parts Il-ll. Parts'lO-ll oooperate with the non-rotative ring il. and the gimbal assembly just described provides for free tilting movement of ring l1 in all directions. A rotative swash ring il is mounted on ring I1 bymeansoibearingsllandtheringtlisprblades, the vertical movement providing for variation of mean pitch oi' all oi' the blades in the scribed more i'ully herebelow.

4). This arm is moved upwardly or downwardly by pitch control mechanism which is described 'below with particular reference to Figures 3, 4,

5 and 8. v

Referring now to Figure 3. a non-rotative sleeve 63 surrounds the rotor supporting sleeve i8 and isvertically movable thereon. Sleeve il same sense, andthe tilting movement effecting a cyclical variation of blade pitch in a sense dei Rotation oi ring 8l with the rotor is assured by means of a scissors linkage 13 (see Figures 3 and 4) which is con,- nected at one end to the ring Il by a universal joint and at the other end by a simple pivot to an arm Il projecting laterally from the rotative hub spindle.

`'Ihe controls for effecting vertical and tilting movements of the swash'ring 6I are shown in Figures 3, 4, 5 and 8. As best seen in' Figures 5, and 8, the control for moving the sleeve Il and thus the ring 88 vertically, includes a lever Il pivoted intermediate its ends on brackets 1l carried by the housing 24, the inner end of the lever being coupled by means oi' links 'Il with an apertured lug projecting from the sleeve 83. The

upper and lower ends of links 11 are connected with the associated parts by joints providing some lost motion in a sense affording limited freedom for angular turning ot the sleeve l) about the axis or the spindle l2. This lost motion is provided to accommodate various control movements described below. The outer end o1' lever 15 is connected with a multi-part link II-19 incorporating complementarily threaded internal and external screw parts, the lower end of link part 19' being connected with a pulley or the like Il which is mounted for rotation in the fixed supporting box 8|. A closed circuit cable I2 is adapted'to actuate the pulley all and thus rotate the lower link part 19, the eiect of which is to causethe `upper link part 18 to move upwardly or downwardly and thus cause the sleeve 63 and the swash ring 8l carried thereby to move `downwardly or upwardly. Cable l2 is adapted to be actuated by the mean pitch control lever 83, the sense of threading between the link parts 18 and 19 and the arrangement of the cable and lever preferably being such vthat rearward movement oi' the lever effects increase o! mean blade pitch and forward movement of the lever elects decrease of mean blade pitch. A control stick Il hooked up in this `sense is illustrated in Figure 8.

I When the sleeve 6I is caused to move upwardly or downwardly by the mean pitch control lever, a similar motion is imparted to the floating hub locking mechanism which includes a sleeve part 84 (see particularly Figures 3 and 7) which is rotatively mounted on the sleeve 63 by means of the bearing 85. The sleevellis provided with three pairs of upstanding elements l-IB provided with laterally projecting stop members 81 which are adapted to engage the nanges -ll :,ssacu.

which project laterally from each of theblade mounting stubs '45. The floating htlocking parts 04 to 01 inclusive, are all rotatable with the rotor and are movable'vertically with the sleeve 83, so that when the mean pitch control lever 03. is adjusted to increase or decrease the mean blade pitch, the-stop members 01 are caused to move upwardly or downwardly. The arrangement of these parts is such that when the mean floating hub, may readily be locked out. In making a landing, as soon as the craft has come to rest on the ground the pilot would normally reduce the mean blade pitch in order to avoid undesired re-take-off, and in effecting this maneuver the floating hub is automatically. locked. Similarly, when a take-oil isA being made, after the rotor has been accelerated to an appropriate take-oi! R. P. M., the pilot would normally move the mean pitch control lever to increase the pitch of the rotor so as to eifect take-off, and this control motion of the mean pitch control lever is accompanied by release of the locking mechanism, so that the rotor system is conditioned for proper flightl operation.

Turning now to the controls 'for the cyclical or differential pitch variation of the rotor blades,

' attention is-firstdirected to Figures 3 and 5 from which it willbe seen that the non-rotative (but vided with a depending part |00 provided 'with a pair of laterally projecting ears v|0||0| to which one end of a-link |02 is connected as by means of a universal joint |03. Motion is im-v partedto link |02 axially thereof bymeans of a bell-crank and multi-part screw threaded link of essentially the same type as described above (reference numerals 0| to Il inclusive); The horizontal arm of this bell-crank appears at |04 in Figure 5 and the upright arm appears at |00, in Figure 3. A universal or ball type joint l|0'| is provided between the link |02 and the upright bell-crank arm |06.` Upward' and downward swinging motion of .bell-crank arm |04 is provided for by the multipart screw threaded link associated therewith in the same general man. ner as described above with reference to the control elements 01, 00 and 80 in Figure 3.

In the case of the screw threaded link device associated with bell-crank lever |04 ,'the rotative motion is provided for by means o f a pulley or sprocket |00 mounted in a ilxed supporting box |09, the pulley or sprocket beingassociated with a cable or chain I0. l

The ball or universal joints (90, |03 and |01- see Figure 5) associated with the bell-crank arms 0| and |00 and with the links 00 and |02, are provided to accommodate the various tilting motions which occur when the swash member is lbeing concurrently tilted on the two axes dened by trunnions 84 and 06. The extension of links 00 and |02 for an appreciable distance in a generally horizontal plane minimizes the eii'ect ofono control on the other. When the v.link |02 is actiltable) ring 61 is provided with a pair of downwardly projecting ears 89-89 to which a yoke member 80 is 'pivotecL The yoke member constitutes alink adapted to be moved to the left or to the right, when viewed as in Figures 3 and 5, so as to tilt the ring 81 and .thus the rotative swash ring 68 about the axis of the tilting trunnions 68. Thismotion is imparted to the yoke or link member 00 by connection of the free end thereof withpne arm 0| of a bell-crank which is.

pivotally mounted on supporting ears 92. the

.otherarm 93 of the bell-crank being connected with a multi-part link 04-95, incorporating cooperating internally and externally threaded elements so that rotation of part 95 causes the part 94 to move upwardly or downwardly. A ball or universal joint 9B is provided between the link 00 and arm 9| of the bell-crank for a purpose to be described hereinafter.

As shown in Figure 3, the Alower end of link .part 95 is associated with a sprocket 91, with place of the pulley-and-cable connection 00-82 illustrated in Figure 5. In the event of employment of a short length of chain adjacent the multi-part threaded link devices, it is contemplated that lengths of cable will be extended therefrom to the controls in the cockpit of the 4 aircraft (see Figure 8 and description herebelow) For the purpose of tilting the non-rotative ring 01 and thus the rotatable swash ring -60 about the axis of trunnions tl-M, the ring 01 is protuated the tilt of the non-'rotative and'rotative swash rings 01 and 00 takes place substantially about the axis of trunnions SL44, the motion of the depending control ear |00 being constrained to occur in asomewhat arcuate path about a center point dened by the ball joint of the other set of control connections. This arcuate motion is accompanied byslight rotation of the gimbal mounting sleeve 03, and such rotation is accommodated by the lost motion connection associated with ends of links 11 of the mean pitch control connections.

In consideringthe orientation of the control hookup providinggfor tilting movements of the swash ring so as to shift the lift line of the rotor laterally and longitudinally for lateral and longitudinal control of the aircraft, it is ilrst pointed out that Figure 8 illustrates both the rotor mount and the control system as viewed from a left, forward position, i. e., looking toward the roto!l mount and controls in a direction approximately diametrically opposite tothe arrow'applied at the lower leftcorner of Figure 4, which arrow, in Figure 4` identifies the Forward direction. With this in mind, it will be seen, Iparticularly from Figure 4. that the mean pitch control levell 10 extends forwardly from the rotor hub substantially in the longitudinal vertical midplane of the aircraft. It will also be seen that the lateral control link |02 and the longitudinal control link l0 extend rearwardly from the rotor mount in directions diverging to the right and to the left, respectively, from the longitudinal vertical midplane (a .plane containing the `arrow labeled i Forward" vin Figure 4) As a result, actuation of the longitudinal control link 90 causes the swash ring 00 to tilt about a horizontal axis intersecting the rotor axis and extended at an angle of approximately 45" from the longitudinal vertical mid-plane, the divergence being to the left of said plane ahead oftherotoraxis-andtotho right of said plane rearwardly of the rotor axis. On the other hand actuation of the lateral control link |02 causes the swash ring 8l to tilt about a horizontal axis intersecting the rotor axis and extended at an angle of` approximately 45 from the longitudinal vertical mid-plane, but with the divergence being to the right of said plane ahead of the rotor axis and to the left of said plane rearwardly of the rotor axis.

With pitch control arms 82 extended (as shown in Figure 4) forwardly from the' blade roots to\a region angularly spaced in advance of the longitudinal blade axes by,l approximately 45.-the above described sense of tilting of the swash ring 88 results in cyclical pitch variation (upon actuation of the lateral or longitudinal controls) in such sense that longitudinal control actuation causes the blades to attain maxi mum and minimum pitch as they pass through advancing and retreating positions located approximately transverselyV of the aircraft, i. e.,

approximately at right angles to the longitudinal .vertical mid-plane. Lateral, control actuation causes the blades to attain maximum and minimum pitch as they pass through positions at the front and rear of the machine, i. e., positions lying substantially in the longitudinal vertical mid-plane ofthe aircraft.

Vas they pass at' the front, and maximum nism of the present invention, it being pointed The cable system for controlling the longitudinal and lateral tilting movement of the swash T ring is illustrated in Figure 8, from which it will be seen that the longitudinal control cable 98 is arranged in a closed circuit associated with the downward projection Ill of the cyclical pitch control leverV H2. the lever being mounted for fore-and-ait tilting movement as by pivots one of which appears at H3. 4The control stick is also mounted for lateral movement as by a pivot ill and the cable IIO is associated with the control stick below the pivot lll in the manner of a closed circuit, so that movement of the control stick to the right or to the left (as indicated by the arrows in Figure 8) actuates the cable H0.

The control cables just described and the lateral and longitudinal pitch Vcontrol levers, links, etc., are so arranged that the points of maximum pitch increase and maximum pitch decrease take place in the senses mentioned just below. Assuming forward ilight of the aircraft, forward movement of the control stick I I2 causes the blades to attain maximum decrease of pitch as they pass on the advancing side of the rotor, and maximum increase of pitch as they pass on the retreating side, substantially diametrically opposite to the position of maximum pitch de-v out that the pitchy control connections and the mounting of .the hub for free iioating to and rotation in diierent planes provides for another type of cyclical pitch variation which takes place automatically during translational flight or during the existence of any other external influence, such Vas an air bump." tending to alter or disturb the mean plane of rotation of the rotor and its hub. The two types of cyclical pitch variation (manually controlled. and automatically set up), are. under various conditions of night, superimposed one upon the other.

'I'he second or automatic type of cyclical pitch variation occurs as a result o'ftiltingfmovement of the floating hub with reference to the plane of rotation of the swash ring 88. When this occurs. for example when the iloating hub tilts downwardly at the forward edge.. the blades experience a cyclical pitch variation, attaining a maximum pitch increase as the blades pass the forward position and a maximum pitch decrease crease. Backward movement of the control stick causes the blades to attain maximum increase of pitch as they pass on the advancing side of the rotor and maximum decrease of pitch as they Apass on the retreating side. The longitudinal control secured in this way is instinctive, resulting in a nose-down moment upon forward movement of the control stick and a nose-up moment upon backward movement of the control stick.

. With respect to lateral control, with a rotor rotating in the direction indicated by the arrow r in Figure 4, movement of the control stick to the right causes the blades to attainv maximum increase or pitch when the blades pass at the front of the machine and maximum decrease of pitch when the blades pass at the rear; andconversely when the control stick is moved to the left the blades attain maximum decrease oi pitch as the blades pass the position at thai-ear of the circle of rotation. Similarly. if the oating hub tilts downwardly at the advancing side `(at the right-hand side of the aircraft. with a direction of rotor rotation such as indicated in Figure 4) a cyclical pitch variation is automatically set up, with the blades attaining maximum pitch increase as they pass on the advancing side of the rotor and a maximum pitch decrease as they pass at the retreating side. 'It may be mentioned that the sense of automatic cyclical pitch varia-l tion here described is defined with relation. Ito a plane of reference fixed with respect to the `aircraft. for instance a plane perpendicular to 'the axis of the non-tilting rotor mounting spindle 32. 'I'he automatic cyclical pitch variation here discussed may alternatively be deilned in terms of maximum pitch increase and maximum pitch decrease with relation to the instantaneous plane of rotation of the oating hub member, in which event the points in the circle of rotation where maximum pitch increase and maximum pitch decrease occur must be differently dellned, i. e., as follows. f

Because of the extension of the pitch control arms 62 forwardly of the blade roots, if the automatic cyclical pitch variation is measured with relation to the instantaneous plane of rotation of the vfloating hub, if the oating hub tilts downwardly at a given azimuth or position in the circle of rotation, the blades will experience maximum increase of pitch as they pass a pointap- .proximately 45 behind (i. e., trailing) the azimuth of downward tilt of the hub; the point of maximum pitch decrease being diametrically opposite.

In any event, regardless of the method of deni nition of the automatic cyclical pitch variation, theeffect of the automatic cyclical pitch variation tends to centralize the mean plane of rotation of the rotor and its hub and to auto- 'matically compensate or correct vable pitch change of the blade.

for disturbing influences. such as air bumps.

Another feature to be noted is here mentioned with reference to Figure 4. As hereinabove described, the drag'pivot pin Il provides freedom for lag-lead movement of the blade. desirably through a range such as that indicated by the lines l-l. It will benoted that considerably greater freedom is provided for in the lagging sense than in the leading sense. the reason for this being that power drive of the rotor through the hub normally causes considerable lagging movement of the blades; whereas, in. autorotative flight, which is also contemplated according to the invention (atleast for purposes of descent without power), the mid-position of the blade with reference to lag-lead displacements approximates a truly radial position. From examination of the positions of the blade pitch control arms B2 and of the swash ring arms 10. it will be noted that when a blade is in truly radial position the free end of arm l! lies somewhat ahead of the free end of arm 1l. Under conditions of power drive lthe blade lags appreciably, and in this condition the tree end of arm 82 shifts rearwardly with respect to the free end of arm 10. With the range of lag-lead motion indicated, the free end of arm 1li is noticeably behind the free end of arm l2 when the blade is in radial (or autorotative) position, but in the average position of power drive, the free end of arm 82 is located substantially vertically above the free end of arm 10, so that in the condition of power drive lag-lead movements of the blade do not appreciably alter the operation of the pitch control. However, when the condition oi' operation changes from that of power drive to that of autorotative night. an appreciably automatic decrease of blade pitch occurs automatically as the blades swing vi'orwardly to the radial position. which is of advantage since. in general. a lower pitch is desiredfor autorotative operation than for power driven operation.

Another advantageous feature of the arrangement disclosed is the location of the free end of the pitch control arm l! close to a position along the axis of the iiapping pivot l2 ifor the blade.

- l2 rangement. moreover. is of advantage in minimising the transmission of undesirable vibrationsor Ishalres to the control stick, and also in minimizv ing the effects of air bumps or other disturbances.

Iclaim: l

1. In an aircraft, a rotor comprising a hub member mounted to rotate about a generally upright axis. radially extending blades each having In Figures 3 and 4. the mean pitch control and the connections extended to the'blades are shown tion (see the left-hand side of Figure 3l the free end of the arm t! is tust a little below the axis of the napping pivot l2. When the pitch is increased to a positive value suitable for flight operation. the free end of arm flies close to or a little above a horizontal plane containing the flapping pivot axis l2.

Because of the location oi' the free end oi' the pitch control arm B2, as just described, flapping movements of the blade about the napping pivt l! take place without introducing any appreci- However. tilting or floating of the hub and rotation thereof in a plane out of parallelism with that of the swash ring 88 automatically introduces a pitch correction tending to bring the plane ot rotation of the hub back into parallelism with the plane'of ro4 tation of the swash ring 8l.

y ticularly in association with the intersecting apping and drag pivots for mounting the blades on the hub, and with the control system, is highly effective in providing smoothness in rotor operaa root end mounting member, mechanism for each blade for pivotally interconnecting the hub member and said mounting member to provide for pivotal movement of the blade as a whole with respect to the hub, said pivot mechanism for each blade including a flapping pivot and a pitch change pivot, controllable means ior positively shifting the blades each as a unit on its pitch change pivot, including a control member rotative with the rotor, means mounting the hub member for tilting movement to and rotation in different planes irrespective of the plane of rotation of said control member, and actuating'connections between the control member and the root end mounting members of the several blades including, for each blade, a control arm extended from the root end mounting member for the blade generally in the rotative path of travel thereof to a position angularly oiiset from the blade substantiallyless than 90', to provide for automatic vari ation of blade pitch upon tilting movement of the hub member. i

2. In an aircraft, a rotor comprising a hub member mounted to rotate about a generally upright axis,radially extending blades each having y planes irrespective of the plane of rotation of said control member. and actuating connections between the control member and the root end mounting members of the several blades including, for each blade, a control arm extended from the root end mounting member for the blade generally in the rotative path of travel thereof Kto provide for automatic variation of blade pitch upon tilting movement of the hub member, in a sense eiecting decrease of rotor blade pitch angle toward that side of the rotor on which the hub member tilts upwardly and increase of the rotor blade .pitch angle toward that side of the rotor on which the hub member tilts downwardly.

I 3. In an aircraft, a sustaining rotor comprising a rotative hub member mounted for tilting to and rotation in different planes, blade means connect ed with the hub memberv and arranged for variation in mean eiiective pitch angle, adiustable control means for regulating the mean pitch angle'of -the blades, and means operated by adjustment of said control means to lock the hub member as against tilting movement.

4. A construction according to claim 3 in which the mean rotor blade pitch angle is adjustable overa range including approximately zero eiiective pitch, and in which the means for locking 'the' hub member, as against tilting movement, is

tion under various conditions of flight.. The arstantially zero pitch setting.

hub with freedom andere `rotation in diilferent planes, blade means con.

nected with the hub member by flapping and pitch change pivots, means associated with the blades and the tiltable hub member for restraining the blades as against excessive downward swinging movement on the flapping pivots. adiustable control means for regulating the mean pitch angle of the blades. and means operated by adjustment of said control means to lock the hub member as against tilting movement.

8.. In an aircraft, a sustaining rotor comprising a rotative hub member mounted for tilting to and rotation in diii'erent planes, blade means connected with the hub member with treedom for flapping movement with respect thereto, means associated with the blades and the tiltable hub member for restraining the blades as against excessive downward swinging movement, and controllable means for locking the hub member as against tilting movement.

7. A construction in laccordance with claim I6 in which the aircraft is further provided with a normal night control organ manipulable by the pilot for maneuvering purposes adJustable toia point beyond normal flying range and means operated by adjustment of said iiight control organ to said point beyond normal flying range to enectvlocking of the hub member as against tilting movement.

8. In an aircraft. a sustaining rotor compris ing a rotative hub. a blade connected with the hub by three pivotal mountings providing. spectively, for lag-lead, flapping and pitch change movement.` the pivotal mounting providing for napping movement being operatively in i4 means mounting the hub for tilting to and rotation in different planes irrespective of the plane of rotation of said member. and actuating connections between the blade and' said ro tatable member providing for cyclical variation oi' blade pitch when the plane of rotation of the hub is angled with respect to the plane of rotation of said rotatable member, said: actuating connections being arranged to effect cyclical blade pitch change in such sense that when the plane of rotation oi' the hub is tilted downwardly toward any angular position in the general circle of rotation ot the rotor the blade experiences,

maximum increase of pitch. with relation to a horizontalplane fixed with respect to the aircraft, when the blade passes through the same azimuthal position as the position of downward tilt of the hub.

10. In an aircraft.a sustaining rotor comprising a rotative hub, a blade connected with the hub with freedom for pitch change movement, mech- 'anism for varying the blade'pitch angle includ ing a member rotatable with the rotor. means mounting the hub for tilting to and rotation in different planes irrespective of the plane of rotation of said member. and actuating connections between the blade and said rotatable mem ber providing ior cyclical variation ofvblade pitch when the plane of rotation of the hub is angled with respect to the,v plane of rotation of said roe tatable membensaid actuatm connections be ins arranged to effect cyclical blade pitch change vin enen sense that -when the plane of rotation ofthe heb is tilted downwardly toward any anguv lar .position in the general circle of rotation oi l the rotor the blade experiences maximum interposed between the other two pivotal mountings, and the pivotal mounting providing for pitch change being located outboard of the other two, mechanism for varying the blade pitch angle including a ,member rotatable with the rotor, means mounting the hub for tilting to andy roe tation in din'erent planes irrespective of the plane ofrotatlon of said member. and actuating connections between the blade and said rotatable member providing for cyclical variation of bla'de pitch when the plane of rotation of the hub is angled with respect to ,the plane of rotation .of said rotatable member. said actuating connections including a flexible `ioint positionedin alignment with the axis of the pivotal mounting providing for blade napping. whereby blade flapping with respect to the hub doesnot apprecia'biy alter the blade pitch established Diem for varying the blade pitch.

lil.l In an aircraft. a sustaining rotor comprising a rotative hub, a blade connected with the y for pitch change movement, mechanism for varying the blade pitch -anglein- Ia member ,rotatable with the `mtos,

by -the meehacrease of pitch with relation to the instantaneous plane of rotation of the hub when the blade passes through a position angularly trailing the .position of downward tilt of the hub, and said actuating connections further being arranged to eiIect said maximum increase ofpitch when lthe blade passes through a position angular-ly Vtraining the position or downward nue tut Ly an angle in the neighborhood of 4l5. y f

nanars s. calunnia anx-namens crrnn The following references are of in the 1file ,of this patent: i

UNITED STATES PATENT! :minibar Name f i te 1.919.089 Breguet et al. July 1l; 1938 `1,986,709 Breguet et al. L dan. 1. .19,35 .2.030.578 ,Flettner .....V 'Feb.` .11, 193i abusos Platt aqu-:3.3. mi

2,811.24?! .Pitcairn Teb. I6, 1948 2,876,523 Bynnestvedt flday 22. 1045 .2,380,580 Cierva July 31. 1945, 2,432.67? Platt ..-v..--.....` Deo. `1,8. m1 

